Combustion chamber/nozzle assembly and fabrication process therefor

ABSTRACT

An integral, lightweight combustion chamber/nozzle assembly for a rocket engine has a refractory metal shell defining a chamber of generally frusto-conical contour. The shell communicates at its larger end with a rocket body, and terminates at its smaller end in a tube of generally cylindrical contour, which is open at its terminus and which serves as a nozzle for the rocket engine. The entire inner surface of the refractory metal shell has a thermal and oxidation barrier layer applied thereto. An ablative silica phenolic insert is bonded to the exposed surface of the thermal and oxidation barrier layer. The ablative phenolic insert provides a chosen inner contour for the combustion chamber and has a taper toward the open terminus of the nozzle. 
     A process for fabricating the integral, lightweight combustion chamber/nozzle assembly is simple and efficient, and results in economy in respect of both resources and time.

ORIGIN OF THE INVENTION

The invention described herein was made by employees of the UnitedStates Government and may be manufactured and used by the Government forgovernmental purposes without the payment of any royalties thereon ortherefor.

CLAIM OF BENEFIT OF PROVISIONAL APPLICATION

Pursuant to 35 U.S.C. §119, the benefit of priority from provisionalapplication No. 60/057,004, with a filing date of Aug. 18, 1997, isclaimed for this non-provisional application.

BACKGROUND OF THE INVENTION

1. Field of the Invention

This invention relates generally to rocket engines. It relates inparticular to a combustion chamber/nozzle assembly for a rocket engineand to a process for its fabrication.

2. Description of Related Art

Conventional combustion/chamber nozzle assemblies for rocket engines areusually actively cooled. That is, they generally contain integralcooling passages for cooling fluid within the combustion chamber andnozzle walls, which tubular cooling passages are fed by manifolds. Acomplex, weighty structure is presented, the fabrication of whichrequires the construction and assemblage of multiple piece parts throughnumerous procedural steps, including machining, plating, welding, andbrazing. Such a complex, weighty structure, as well as its complicatedmethod of fabrication, are both disadvantageous and in need ofimprovement, as is well known in this art.

SUMMARY OF THE INVENTION

It is accordingly a primary object of the present invention to providewhat is lacking in the prior art, especially a simple, yet highlyefficient, lightweight integral combustion chamber/nozzle assembly for arocket engine. It is also a primary object of the present invention toprovide an uncomplicated and highly reliable process for the fabricationof a simple, lightweight integral combustion chamber/nozzle assembly fora rocket engine, which process is effective and highly efficient,especially in respect of the utilization of time and materials.

These objects and other related benefits are achieved by the presentinvention, which in one aspect thereof is an integral, lightweightcombustion chamber/nozzle assembly which has a shell of a refractorymaterial, such as an alloy of niobium, having a configuration defining achamber of generally frusto-conical contour. The chamber communicates atits larger end with a rocket body, and terminates at its smaller end ina tube open at its terminus, which serves as a nozzle for the rocketengine. The inner surface of the chamber has applied thereto a thermaland oxidation barrier layer, especially of a silicide or aluminum oxide.An ablative silica phenolic insert, which is bonded to the thermal andoxidation barrier layer, is configured to provide a chosen inner contourfor the combustion chamber.

The ablative silica phenolic insert additionally has a taper orreduction in thickness toward the open terminus of the nozzle.

In another aspect, the present invention is a process for fabricatingthe integral, lightweight combustion chamber/nozzle assembly, whichprocess is set forth in detail hereinafter.

The integral, combustion chamber/nozzle assembly according to thepresent invention is simple in design and is much lighter thanconventional combustion chamber/nozzle assemblies, making it a highlydesirable replacement assembly for these reasons alone. Moreover, theintegral combustion chamber/nozzle assembly according to the presentinvention is not actively cooled, as are the assemblies of the priorart, so that there is no need for cooling passages therein. Fabricationis therefore greatly simplified, and accordingly accelerated, resultingin a highly desirable economy in respect of both resources and time.

During the firing operation of a rocket engine employing the integral,lightweight combustion chamber/nozzle assembly of the present invention,resins boil off from the ablative silica phenolic insert, therebycooling the inner surface of the refractory metal shell and leavingbehind a layer of char. This layer of char, along with the remainingsilica phenolic layer, acts as an insulator and protects the refractorymetal shell against overheating. The thickness of the ablative insert ischosen so that the layer of char does not penetrate too deeply duringthe design life of the combustion chamber/nozzle assembly.

During the firing operation of a rocket engine employing the integral,lightweight combustion chamber/nozzle assembly of the present invention,the temperature inside the combustion chamber/nozzle assembly decreasestoward the open end of the nozzle. Therefore the thickness of theablative silica phenolic insert is fashioned to taper down toward theopen terminus of the nozzle, at which terminus the refractory metal can,along with the thermal and oxidation layer applied thereto, survivewithout any ablative protection.

BRIEF DESCRIPTION OF THE DRAWINGS

For a more complete understanding of the present invention, includingits primary objects and attending benefits, reference should be made tothe Detailed Description of the Preferred Embodiments, which is setforth below. This description should be read in conjunction with theattached drawings, wherein:

FIG. 1 is a schematic depiction of an integral, lightweight combustionchamber/nozzle assembly for a rocket engine, according to the presentinvention, showing the cooperative combination of its essentialcomponents;

FIG. 2 schematically depicts a refractory metal shell sprayed onto agraphite mandrel by vacuum plasma spraying, according the process of thepresent invention; and

FIG. 3 schematically depicts a vacuum plasma sprayed shell of arefractory metal deposited on a graphite mandrel with additionalmaterial deposited so that a flange can be machined after the mandrelhas been removed.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

Referring now to the drawings, FIG. 1 schematically depicts theintegral, lightweight combustion chamber/nozzle assembly for a rocketengine according to the present invention as including a refractorymetal shell 10 which defines a chamber of generally frusto-conicalcontour. Refractory metal shell 10 communicates at its base or largerend with a rocket body, and terminates at its smaller end in a tube ofgenerally cylindrical contour which is open at the end thereof and whichserves as a nozzle for the rocket engine. Projecting rim or integralflange 11 is formed at the open end of the nozzle. The refractory metalshell 10 has a protective inner layer 12 which acts as an oxidation aswell as a thermal barrier. Bonded inside the shell is an ablative silicaphenolic insert 13.

The shell 10 is manufactured from a refractory niobium alloy such asNiobium (Columbium) C-129Y, which is available commercially. The methodof manufacture consists of vacuum plasma spraying (VPS) the alloy 10 ona graphite mandrel 14, as seen in FIG. 2, using standard techniques,well known in the art. The inner protective layer 12 may be applied byeither VPS or by coating. The VPS method applies a layer such asaluminum oxide to the graphite mandrel and then transitions into therefractory alloy. Using this option, VPS is started-initially with thethermal and oxidation barrier material such as aluminum oxide, and thena gradual transition is made to the refractory metal, followed bycontinuing with the refractory metal to provide the required thicknessthereof. An alternative method is to apply a silicide coating bystandard methods to the shell interior after it has been removed fromthe mandrel. The difference between the coefficients of thermalexpansion of the refractory metal and the graphite permits the shell tobe easily removed from the mandrel after the parts cool following theVPS process. FIG. 3 shows a vacuum plasma sprayed shell 10 deposited onthe graphite mandrel 14 with additional material 15 deposited on thesmall end so a flange can be machined after the mandrel is removed.Another method of fabricating the flange is to machine it from wroughtmaterial and then weld it onto the VPS deposited shell.

The ablative silica phenolic insert, 12, is made by wrappingcommercially available silica phenolic tape on a steel mandrel which isconfigured to produced the desired inner contour of the combustionchamber. The tape is laid up at an angle to the part centerline, so thatthe edge of the tape will be exposed to the high internal temperatures.This tape wrapped billet is cured using standard techniques and then theexterior is machined to match the interior of the refractory metalshell. The ablative insert is then bonded onto the inside of the metalshell, completing the combustion chamber/nozzle assembly.

The silica phenolic insert takes advantage of an ablative process.During a firing of the rocket engine, resins boil off from the silicaphenolic, cooling the surface of the refractory metal shell and forminga char layer. This layer acts as an insulator, protecting the refractorymetal shell. The ablative thickness is chosen such that the char layerwill not penetrate too deeply during the design life of the unit. As thetemperature drops toward the exit of the nozzle, the ablative insert isextended to a point where the temperature is low enough to be handled bythe refractory metal in combination with its protective inner layer orcoating.

If the temperature of the interface between the ablative insert and therefractory metal shell is too high to achieve a reliable bond,mechanical attachment, such as pinning, can be incorporated.

We claim:
 1. An integral, lightweight combustion chamber/nozzle assemblyfor a rocket engine, which assembly comprises:a shell of a refractorymetal having a configuration defining a chamber of generallyfrusto-conical contour, the refractory metal shell having an inner andan outer surface and communicating at its base or larger end with arocket body, and terminating at its smaller end in a tube of generallycylindrical contour which is open at the terminus thereof and whichserves as a nozzle for the rocket engine; the entire inner surface ofthe refractory metal shell having applied thereto a thermal andoxidation barrier layer; and an ablative silica phenolic insert bondedto the exposed surface of the thermal and oxidation barrier layer andconfigured to provide an inner contour for the combustion chamber, theablative silica phenolic insert having a taper or gradual reduction inthickness thereof toward the open terminus of the nozzle.
 2. Theintegral, lightweight combustion chamber/nozzle assembly of claim 1,wherein the refractory metal is an alloy of niobium.
 3. The integral,lightweight combustion chamber/nozzle assembly of claim 2, wherein thethermal and oxidation barrier layer is a silicide.
 4. The integral,lightweight combustion chamber/nozzle assembly of claim 2, wherein thethermal and oxidation barrier layer is aluminum oxide.